Geared turbomachine architecture having a low profile core flow path contour

ABSTRACT

An exemplary geared turbomachine assembly includes a core inlet having a radially inner boundary that is spaced a first radial distance from a rotational axis of a turbomachine, and a compressor section inlet having a radially inner boundary that is spaced a second radial distance from the rotational axis. A ratio of the second radial distance to the first radial distance is of about 0.65 to about 0.9.

CROSS-REFERENCE TO RELATED APPLICATIONS

This disclosure claims priority to U.S. Provisional Application No.61/593230, filed Jan. 31, 2012, and is incorporated herein by reference.

BACKGROUND

This disclosure relates to a geared turbomachine having a core inletradially spaced from a compressor inlet.

Turbomachines, such as gas turbine engines, typically include a fansection, a turbine section, a compressor section, and a combustorsection. The fan section drives air along a core flow path into thecompressor section. The compressed air is mixed with fuel and combustedin the combustor section. The products of combustion are expanded in theturbine section.

A core inlet controls flow of air into the core flow path. The flow ofair moves from the core inlet to a compressor section inlet. Therelative radial positions of the core inlet and the compressor sectioninlet influence flow through the core and a profile of the turbomachine.

SUMMARY

A turbomachine assembly having a geared architecture according to anexemplary aspect of the present disclosure includes, among other things,a core inlet having a radially inner boundary that is spaced a firstradial distance from a rotational axis of a turbomachine, and acompressor section inlet having a radially inner boundary that is spaceda second radial distance from the rotational axis. The ratio of thesecond radial distance to the first radial distance is of about 0.65 toabout 0.9.

In a further non-limiting embodiment of the foregoing turbomachineassembly having a geared architecture, the assembly may include aradially inner boundary of the core inlet at the location of a coreinlet stator.

In a further non-limiting embodiment of either the foregoingturbomachine assemblies having a geared architecture, the assembly mayinclude a radially inner boundary of the compressor section inlet at thelocation of a compressor rotor.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies having a geared architecture, the assembly mayinclude a compressor rotor that is a first stage rotor of a low-pressurecompressor.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies having a geared architecture, the core inlet maybe an inlet to a core section of the turbomachine.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies having a geared architecture, the assembly mayhave an inlet flow of the compressor section that is from about 30lb/sec/ft² to 37 lb/sec/ft² when the turbomachine is operating at acruise speed.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies having a geared architecture, the assembly mayhave a turbine inlet temperature of a high-pressure turbine within theturbomachine that is from about 2,000 F to 2,600 F when the turbomachineis operating at a cruise speed.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies having a geared architecture, the assembly mayhave a blade in the compressor section having a tip speed duringoperation that is from about 1,050 fps to 1,350 fps.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies having a geared architecture, the assembly mayinclude a fan section of the turbomachine driven by a gearedarchitecture that is driven by a shaft that rotates a compressor rotorwithin the compressor section.

In a further non-limiting embodiment of any of the foregoingturbomachine assemblies having a geared architecture, the assembly mayinclude a geared architecture that has a gear reduction ratio from about2.2 to 4.

A gas turbine engine assembly according to an exemplary aspect of thepresent disclosure includes, among other things, a core inlet statorhaving a stator root that is spaced a first radial distance from arotational axis of a gas turbine engine, and a compressor blade within afirst stage of a compressor section. The compressor blade has a bladeroot that is spaced a second radial distance from the rotational axis. Aratio of the second radial distance to the first radial distance is ofabout 0.65 to about 0.9.

In a further non-limiting embodiment of the foregoing gas turbine engineassembly, the assembly may include a stator root radially aligned with aradially inner boundary of a core flow path through the gas turbineengine.

In a further non-limiting embodiment of either of the foregoing gasturbine engine assemblies, the assembly may include a blade rootradially aligned with a radially inner boundary of a core flow paththrough the turbine engine.

In a further non-limiting embodiment of any of the foregoing gas turbineengine assemblies, the assembly may include a core inlet statorpositioned within an inlet to a core section of the gas turbine engine.

In a further non-limiting embodiment of any of the foregoing gas turbineengine assemblies, the assembly may have an inlet flow of the compressorsection that is from 30 lb/sec/ft² to 37 lb/sec/ft² when the gas turbineengine is operating at a cruise speed.

In a further non-limiting embodiment of any of the foregoing gas turbineengine assemblies, the assembly may have a turbine inlet temperature ofa high-pressure turbine within the gas turbine engine that is from 2,000F to 2,600 F when the gas turbine engine is operating at a cruise speed.

In a further non-limiting embodiment of any of the foregoing gas turbineengine assemblies, the assembly may include the compressor blade havingtip speed during operation that is from 1,050 fps to 1,350 fps when thegas turbine engine is operating at cruise speed.

In a further non-limiting embodiment of any of the foregoing gas turbineengine assemblies, the assembly may include a fan section of theturbomachine driven by a geared architecture that is driven by a shaft.The geared architecture having a gear reduction ratio from 2.2 to 4.

A compressor module of a gas turbine engine according to an exemplaryaspect of the present disclosure includes, among other things, acompressor module a first inner boundary facing radially outward anddefining in-part a core inlet, the first inner boundary located at afirst radial distance from the rotational axis. The compressor moduleincludes a second inner boundary facing radially outward and locatedaxially downstream of the first inner boundary. The second innerboundary is located at a second radial distance from the rotationalaxis. A ratio of the second radial distance to the first radial distanceis of about 0.65 to about 0.9.

In a further non-limiting embodiment of the foregoing compressor module,the radially inner boundary of the compressor section inlet is at alocation of a compressor rotor and the radially inner boundary of thecore inlet is at a location of a core inlet.

DESCRIPTION OF THE FIGURES

The various features and advantages of the disclosed examples willbecome apparent to those skilled in the art from the detaileddescription. The figures that accompany the detailed description can bebriefly described as follows:

FIG. 1 shows a partial section view of an example turbomachine.

FIG. 2 shows a close-up view of a core inlet portion of the FIG. 1turbomachine.

DETAILED DESCRIPTION

Referring to FIG. 1, an example turbomachine, which is a two-spool gasturbine engine 20, generally includes a fan section 22, a compressorsection or module 24, a combustion section 26, and a turbine section 28.Other examples may include an augmentor section (not shown) among othersystems or features.

Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with turbofans. Thatis, the teachings may be applied to other types of turbomachines andturbine engines including three-spool architectures.

In the example engine 20, air moves from the fan section 22 to a bypassflow path B or a core flow path C. The core flow path C is within a coresection of the engine 20. Within the core flow path, compressed air fromthe compressor section 24 communicates through the combustion section26. The products of combustion expand through the turbine section 28.

The example engine 20 generally includes a low-speed spool 30 and ahigh-speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36. Thelow-speed spool 30 and the high-speed spool 32 are rotatably supportedby several bearing systems 38. It should be understood that variousbearing systems 38 at various locations may alternatively, oradditionally, be provided.

The low-speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low-pressure compressor 44, and a low-pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than thelow-speed spool 30.

The high-speed spool 32 includes an outer shaft 50 that interconnects ahigh-pressure compressor 52 and high-pressure turbine 54.

The combustion section 26 includes a circumferentially distributed arrayof combustors 56 generally arranged axially between the high-pressurecompressor 52 and the high-pressure turbine 54.

The inner shaft 40 and the outer shaft 50 are concentric and rotate viabearing systems 38 about the engine central longitudinal axis A, whichis collinear with the longitudinal axes of the inner shaft 40 and theouter shaft 50.

A mid-turbine frame 58 of the engine static structure 36 is generallyarranged axially between the high-pressure turbine 54 and thelow-pressure turbine 46. The mid-turbine frame 58 supports bearingsystems 38 in the turbine section 28.

In the example engine 20, the core airflow C is compressed by thelow-pressure compressor 44 then the high-pressure compressor 52, mixedand burned with fuel in the combustors 56, then expanded over thehigh-pressure turbine 54 and low-pressure turbine 46. The high-pressureturbine 54 and the low-pressure turbine 46 rotatably drive therespective high-speed spool 32 and low-speed spool 30 in response to theexpansion. The high-pressure turbine 54 and the low-pressure turbine 46drive a compressor rotor of the high-pressure compressor 52 and acompressor rotor of the low-pressure compressor 44, respectively.

In some non-limiting examples, the engine 20 is a high-bypass gearedaircraft engine. In a further example, the engine 20 bypass ratio isgreater than about six (6:1).

The geared architecture 48 of the example engine 20 includes anepicyclic gear train, such as a planetary, star, or differential gearsystem. The example epicyclic gear train has a gear reduction ratio ofgreater than about 2.2 (i.e., 2.2:1). In some examples, the gearreduction ratio is from about 2.2 to 4. In such examples, theseparameters result in superior engine fuel consumption characteristics.These ranges of ratios help define the inner diameter of the core flowpath in these examples. The low-pressure turbine 46 pressure ratio ispressure measured prior to inlet of low-pressure turbine 46 as relatedto the pressure at the outlet of the low-pressure turbine 46 prior to anexhaust nozzle of the engine 20. In one non-limiting embodiment, thebypass ratio of the engine 20 is greater than about ten (10:1), the fandiameter is significantly larger than that of the low pressurecompressor 44, and the low-pressure turbine 46 has a pressure ratio thatis greater than about 5 (5:1). In prior art designs, significantlylarger sized, lower-speed turbines are required to achieve this samerange of ratios. In this embodiment, the low-pressure turbine is smallerand runs faster than the prior art designs. The geared architecture 48of such embodiments is an epicyclic gear train with a gear reductionratio of greater than about 2.4 (i.e., 2.4:1). It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present disclosure isapplicable to other gas turbine engines including direct driveturbofans.

In the example engine 20, a significant amount of thrust is provided bythe bypass flow B due to the high bypass ratio. The fan section 22 ofthe engine 20 is designed for a particular flight condition—typicallycruise at about 0.8 Mach and about 35,000 feet. This flight condition,with the engine 20 at its best fuel consumption, is also known as bucketcruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industrystandard parameter of fuel consumption per unit of thrust. Althoughdescribed above as about 0.8 Mach, cruise may range from about 0.7 to0.9 Mach.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. The low FanPressure Ratio according to one non-limiting embodiment of the exampleengine 20 is less than 1.45. This ratio is a characteristic of many ofthe disclosed examples.

Low Corrected Fan Tip Speed is the actual fan tip speed divided by anindustry standard temperature correction of “T”/518.7^(0.5). Trepresents the ambient temperature in degrees Rankine. The Low CorrectedFan Tip Speed according to one non-limiting embodiment of the exampleengine 20 is less than about 1150 fps (351 m/s). Low fan tip speed areoften desirable for good fan efficiency and low noise.

In some examples, the tip speed of the low-pressure compressor 44 isfrom about 1,050 fps (320 m/s) to 1350 fps (411 m/s). Speeds within thisrange may balance fuel consumption, with low stage number and partscount.

Referring now to FIG. 2 with continued reference to FIG. 1, the coreflow path of the example engine 20 begins at a core inlet 60 and extendsthrough and past the low-pressure compressor 44. The core inlet 60 has aradially inner boundary 62 and a radially outer boundary 66.

A core inlet stator 70 is located at or near the core inlet 60. The coreinlet stator 70 attaches to a core case 74 at the radially innerboundary 62. The core inlet stator 70 attaches to an inlet case 78 atthe radially outer boundary 66. The core inlet stator 70 extendsradially across the core flow path C.

In this example, the radially inner boundary 62 is positioned a radialdistance D₁ from the axis A. The distance Di, in this example, alsocorresponds to the radial distance between a root 64 of the core inletstator 70 and the axis A. In this example, the root 64 of the core inletstator 70 is radially aligned with the radially inner boundary 62 of thecore flow path C.

After flow moves through the core inlet 60, the flow moves through acompressor inlet 82 into the compressor section 24. In this example, thecompressor section inlet 82 is an inlet to the low-pressure compressor44 of the compressor section 24. The compressor inlet 82 extends from aradially inner boundary 86 to a radially outer boundary 90.

Notably, a blade 98 of a rotor within the low-pressure compressor 44extends from a root 102 to a tip 106. The blade 98 is located at or nearthe compressor inlet 82. The blade 98 part of a compressor rotor withina first stage of the compressor section 24. The blade 98 is thus part ofa first stage rotor, or a leading blade of the compressor section 24relative to a direction of flow along the core flow path C.

In some examples, the blade 98 represents the axial position where airenters the compressor section 24 of the core flow path C. The blade 98extends radially across the core flow path C.

The radially inner boundary 86 is positioned a radial distance D₂ fromthe axis A. The distance D₂, in this example, also corresponds to theradial distance between the root 102 of the blade 98 and the axis A. Inthis example, the root 102 is radially aligned with the radially innerboundary 86 of the core flow path C.

In the example engine 20, a preferred ratio range of the distance D₂ tothe distance D₁ spans from about 0.65 to about 0.9, which provides arelatively low profile core flow path contour. High profile flow pathcontours have greater differences between D₂ and D₁, and thus larger“humps” between the core inlet 60 and the compressor inlet 82. Highprofile flow path contours introduce discontinuities that undesirablydisrupt the airflow and undesirably add weight to the engine 20. Theratio range of about 0.65 to 0.9 is made possible, in part, by theincorporation of the geared architecture 48 into the engine 20. The“hump” in this example is generally area 100.

Other characteristics of the engine having this ratio may include theengine 20 having a specific inlet flow of the low pressure compressor atcruising speeds to be between 30 lb/sec/ft² to 37 lb/sec/ft². Thespecific inlet flow is the amount of flow moving into the compressorsection 24 and specifically, in this example, into a compressor inlet 82and through the compressor section 24.

Another characteristic of the example engine 20 is that the cruisespeeds of the example engine are generally Mach numbers of about 0.7 to0.9.

Yet another characteristic of the engine 20 is that a temperature at aninlet to the high-pressure turbine 54 may be from 2,000° F. (1093.33°C.) to 2,600° F. (1426.66° C.). Maintaining temperatures within thisrange balance of good fuel consumption, low engine weight, and lowengine maintenance costs.

Yet another characteristic of the engine 20 is that a tip speed ofblades in a rotor of the low-pressure compressor 44 (a compressor rotor)may be from about 1,050 fps (320 m/s) to 1,350 fps (411 m/s).

In this example, the geared architecture 48 of the engine 20 may have agear ratio of from about 2.2 to 4.

The preceding description is exemplary rather than limiting in nature.Variations and modifications to the disclosed examples may becomeapparent to those skilled in the art that do not necessarily depart fromthe essence of this disclosure. Thus, the scope of legal protectiongiven to this disclosure can only be determined by studying thefollowing claims.

We claim:
 1. A turbomachine having a geared architecture, comprising: acore inlet having a radially inner boundary that is spaced a firstradial distance from a rotational axis of a turbomachine assembly; and acompressor section inlet having a radially inner boundary that is spaceda second radial distance from the rotational axis, wherein a ratio ofthe second radial distance to the first radial distance is of about 0.65to about 0.9.
 2. The turbomachine having a geared architecture of claim1, wherein the radially inner boundary of the core inlet is at alocation of a core inlet stator.
 3. The turbomachine having a gearedarchitecture of claim 1, wherein the radially inner boundary of thecompressor section inlet is at a location of a compressor rotor.
 4. Theturbomachine having a geared architecture of claim 3, wherein thecompressor rotor is a first stage rotor of a low-pressure compressor. 5.The turbomachine having a geared architecture of claim 1, wherein thecore inlet is an inlet to a core section of the turbomachine.
 6. Theturbomachine having a geared architecture of claim 1, wherein an inletflow of the compressor section is from about 30 lb/sec/ft² to 37lb/sec/ft² when the turbomachine is operating at a cruise speed.
 7. Theturbomachine having a geared architecture of claim 1, wherein a turbineinlet temperature of a high-pressure turbine within the turbomachine isfrom about 2,000 F to 2,600 F when the turbomachine is operating at acruise speed.
 8. The turbomachine having a geared architecture of claim1, wherein a tip speed of a blade array in the compressor section duringoperation is from about 1,050 fps to 1,350 fps.
 9. The turbomachinehaving a geared architecture of claim 1, wherein a fan section of theturbomachine is driven by a geared architecture that is driven by ashaft that rotates a compressor rotor within the compressor section. 10.The turbomachine having a geared architecture of claim 9, wherein thegeared architecture has a gear reduction ratio from about 2.2 to
 4. 11.A gas turbine engine, comprising: a core inlet stator having a statorroot that is spaced a first radial distance from a rotational axis of agas turbine engine; and a compressor blade within a first stage of acompressor section, the compressor blade having a blade root that isspaced a second radial distance from the rotational axis, wherein aratio of the second radial distance to the first radial distance is ofabout 0.65 to about 0.9.
 12. The gas turbine engine assembly of claim11, wherein the stator root is radially aligned with a radially innerboundary of a core flow path through the gas turbine engine.
 13. The gasturbine engine assembly of claim 11, wherein the blade root is radiallyaligned with a radially inner boundary of a core flow path through theturbine engine.
 14. The gas turbine engine assembly of claim 11, whereinthe core inlet stator is positioned within an inlet to a core section ofthe gas turbine engine.
 15. The gas turbine engine assembly of claim 11,wherein an inlet flow of the compressor section is from 30 lb/sec/ft² to37 lb/sec/ft² when the gas turbine engine is operating at a cruisespeed.
 16. The gas turbine engine assembly of claim 11, wherein aturbine inlet temperature of a turbine within the gas turbine engine isfrom 2,000 F to 2,600 F of when the gas turbine engine is operating at acruise speed.
 17. The gas turbine engine of claim 11, wherein a tipspeed the compressor blade during operation is from 1,050 fps to 1,350fps when the gas turbine engine is operating at cruise speed.
 18. Thegas turbine engine of claim 11, wherein a fan section of theturbomachine is driven by a geared architecture that is driven by ashaft, the geared architecture having a gear reduction ratio from 2.2 to4.
 19. A compressor module of a gas turbine engine having a rotationalaxis, the compressor module comprising: a first inner boundary facingradially outward and defining in-part a core inlet, the first innerboundary located at a first radial distance from the rotational axis;and a second inner boundary facing radially outward and located axiallydownstream of the first inner boundary, the second inner boundarylocated at a second radial distance from the rotational axis, wherein aratio of the second radial distance to the first radial distance is ofabout 0.65 to about 0.9.
 20. The compressor module of claim 19 whereinthe second inner boundary defines in-part a compressor inlet of thecompressor module.